Cambered aero-engine inlet

ABSTRACT

An inlet of a nacelle for channeling inlet airflow to a fan of a gas turbine engine includes a forward lip and an inner surface. The forward lip extends around a centerline axis of the engine and is adapted to accommodate a flow angle of incoming airflow. The inner surface extends from the forward lip along the centerline axis of the engine to adjacent the fan. The inner surface has a first profile above the centerline axis of the engine and a second profile below the centerline axis of the engine. The first profile and the second profile are adapted to define an inlet centerline axis that extends below the centerline axis of the engine at a face of the fan.

BACKGROUND

The present invention relates generally to gas turbine engines, and moreparticularly, to turbofan aircraft engines.

One of the primary design criteria for aircraft turbofan engines is topropel an aircraft in flight with maximum efficiency, thereby reducingfuel consumption. Thus, turbofan engines are continually being developedand improved to maximize thrust capability with the greatest aerodynamicefficiency possible.

One of the components of the turbofan engine is an array of fan blades,which are positioned adjacent the forward portion of the turbofanadjacent the turbofan's inlet. The fan blades produce thrust, and thus,are typically designed to maximize the aerodynamic loading and theamount of propulsion thrust generated thereby during operation. However,fan loading is limited by stall, flutter, or other instabilityparameters of the air being pressurized.

Fan stall margin is a fundamental design requirement for the turbofanand is affected by aerodynamic fan loading. A major factor affecting theaerodynamic loading of the fan is the geometry of the inlet upstream ofthe fan. Aircraft wings are known to induce an upwash velocity inairflow in the inlet. Unfortunately, conventional inlets do not accountsufficiently for the vertical component (vector) of the airflow. Thecircumferential vector of the airflow causes the fan to operate withdifferent circumferential sectors having different flow-pressure ratiocharacteristics. Specifically, the circumferential vector of the inletairflow manifests itself as a swirl velocity along the face of the fanblades. The swirl velocity is in a direction counter to the direction ofrotation of the fan along at least a portion of the fan face. The swirlvelocity has a destabilizing effect on the flow condition over sector(s)of the fan, and thus, degrades the stall margin of the fan.

Similarly, the different flow-pressure ratio characteristics on the fanproduce a circumferential variation in total pressure at the inlet tothe core stream of the turbofan. The total pressure variation(distortion) has a destabilizing effect on the operation of the lowpressure compressor and high pressure compressor sections of theturbofan.

SUMMARY

An inlet of a nacelle for channeling inlet airflow to a fan of a gasturbine engine includes a forward lip and an inner surface. The forwardlip extends around a centerline axis of the engine and is adapted toaccommodate a flow angle of incoming airflow. The inner surface extendsfrom the forward lip along the centerline axis of the engine to adjacentthe fan. The inner surface has a first profile above the centerline axisof the engine and a second profile below the centerline axis of theengine. The first profile and the second profile are adapted to definean inlet centerline axis that extends below the centerline axis of theengine at a face of the fan.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a schematic representation of a portion of an aircraft havinga wing mounted gas turbine engine including a nacelle inlet.

FIG. 2 is a partially schematic partially sectional view of the turbofanportion of the gas turbine engine of FIG. 1.

FIG. 3 is a schematic sectional view of the nacelle inlet and a fanshown in FIG. 2.

FIG. 4 is a diagrammatic view of the front of the fan showing a decreasein an airflow swirl velocity.

DETAILED DESCRIPTION

FIG. 1 shows an exemplary portion of an aircraft 10 having at least onegas turbine engine 12 mounted to one wing 14 thereof. The engine 12includes a nacelle 16 that defines an inlet 18.

The engine 12 is disposed around an engine centerline axis 20. Thenacelle 16 is an aerodynamic structure that surrounds at least a portionof the engine 12 and forms an inlet 18 thereto. The nacelle 16 channelsfree stream airflow 22 through the inlet 18 to the internal componentsof the engine 12. Adjacent to the inlet 18 and wing 14, the free streamairflow 22 has an upwash angle β with respect to the engine centerlineaxis 20. The upwash angle β is due to the aerodynamic affects of thewing 14. The upwash angle β exists in the inlet airflow after the freestream airflow 22 enters the inlet 18.

As will be discussed subsequently, the portion of the nacelle 16 thatdefines the inlet 18 is adapted to align the upwash angle β with theengine centerline axis 20 at a fan within the engine 12. The alignmentof the upwash angle β with the engine centerline axis 20 reduces oreliminates the vertical vector of the inlet airflow that causes a swirlvelocity along the face of the fan blades. The reduction or eliminationof the swirl velocity allows components of the engine 12 to better meetfan stall margin requirements and improves aerodynamic fan loading.

FIG. 2 shows a section of the turbofan portion of the gas turbine engine12. The engine 12 includes a fan case 24 having a fan case flange 26, afan 28, a low pressure compressor 30, a high pressure compressor 32, acombustor 34, a high pressure turbine 36, and a low pressure turbine 38.The nacelle 16 includes forward lip portion 40, a crown portion 42, akeel portion 44 and an inner surface 46.

In FIG. 2, a section of the nacelle 16 has been removed along a verticalplane extending through top dead center and bottom dead center thereof.The nacelle 16 connects to and extends around the axisymmetrical fancase 24. The fan case flange 26 extends generally radially outward fromthe forward portion of the fan case 24 to connect to the nacelle 16. Theportion of the nacelle 16 that defines the inlet 18 extends forward ofthe fan case 24 and fan case flange 26. The fan 28 is rotationallymounted within the fan case 24 and is co-aligned along the same axis asthe engine centerline axis 20. The low pressure compressor 30, highpressure compressor 32, combustor 34, high pressure turbine 36, and lowpressure turbine 38 extend in series axially downstream (rearward) ofthe fan 28.

The forward lip 40 of the nacelle 16 defines the forward portion of theinlet 18. The forward lip 40 extends around the engine centerline axis20. The nacelle 16 can generally be divided into the crown portion 42,which extends generally above the engine centerline axis 20, and thekeel portion 44, which extends generally below the engine centerlineaxis 20. The crown portion 42 is asymmetrical with respect to the keelportion 44 but both together integrally extend around the enginecenterline axis 20. The keel 44 can be dropped (i.e. adapted to have itsportion of the forward lip 40 be disposed to the rearward of the forwardlip 40 of the crown portion 42) such that the forward lip 40 does notextend along a vertical plane. This arrangement allows the inlet 18 toaccommodate the upwash angle β of incoming airflow.

The inner surface 46 of the nacelle 16 defines a portion of the inlet 18which also extends through the annular fan case 24 to the fan caseflange 26. As will be discussed subsequently, the inner surface 46 ofthe nacelle 16 is profiled on both the crown portion 42 and the keelportion 44 to align the upwash of inlet airflow with the enginecenterline axis 20 at the face of the fan 28.

During operation of the engine 12, the free stream airflow 22 enters theinlet 18, is aligned by the geometry of the nacelle 16, and ispressurized by the fan 28. A portion of the airflow is channeled toengine core where it is pressurized, mixed with fuel/combusted, andexpanded before being expelled from the engine 12. A second portion ofthe airflow bypasses the engine core and is expelled from an outletportion of the nacelle 16.

After being pressurized in the low pressure compressor 30 and highpressure compressor 32, the air is mixed with fuel in the combustor 34for generating hot combustion gases. From the combustor 32, the gasesand airflow are discharged downstream into the high pressure turbine 34.The high pressure turbine 34 and low pressure turbine 36 in turn receivethe combusted gases and extract energy therefrom. The high pressureturbine 34 is joined by a rotor or shaft (not shown) to the highpressure compressor 32 and the fan 28 for powering these componentsduring operation. The combination of thrust produced from the fan 28 andthe components of the engine core propel the aircraft 10 in flight (FIG.1).

FIG. 3 shows a schematic sectional view of one embodiment of the nacelle16 and the fan 28. The axially forward most extent of the fan caseflange 26 is denoted by a fan case flange plane 48. Similarly, the axialforward most portion of the fan 28 is indicated by a fan face plane 50.The inner surface 46 of the nacelle 16 includes a first profile 52 and asecond profile 54. The first profile 52 has a generally convex forwardsection 52F that transitions to a generally concave rearward section 52R(when observed from the engine centerline axis 20) at inflection point53. Similarly, the second profile 54 has a generally convex forwardsection 54F that transitions to a generally concave rearward section 52R(when viewed from the engine centerline axis 20) at inflection point 55.The geometry of the inner surface 46 of the nacelle 16 defines an inletcenterline axis 56.

The forward lip 40 of the nacelle 16 defines the orifice of the inlet18. The forward lip 40 extends around the engine centerline axis 20. Thecrown portion 42 of the nacelle 16 extends above the engine centerlineaxis 20. The keel portion 44 of the nacelle 16 extends below the enginecenterline axis 20. Together both the crown portion 42 and the keelportion 44 form the inner surface 46. The inner surface 46 defines theinlet 18, which extends axially along the engine centerline axis 20 fromthe forward lip 40 to the fan case flange plane 48. The inner surface 46also defines the inlet centerline axis 56 which approximates thegeometric center of inlet 18.

The shape of the inner surface 46 above the engine centerline axis 20 isasymmetrical with respect to the shape of the inner surface 46 below theengine centerline axis 20. More particularly, the inner surface 46 abovethe engine centerline axis 20 is shaped with the first profile 52, whichextends generally axially along the engine centerline axis 20 from theforward lip 40 to at least the fan case flange plane 48. The firstprofile 52 varies in the radial distance it is disposed from the enginecenterline axis 20. Similarly, the inner surface 46 below the enginecenterline axis 20 is shaped with the second profile 54, which extendsgenerally axially along the engine centerline axis 20 from the forwardlip 40 to at least the fan case flange plane 48. The second profile 54varies in the radial distance it is disposed from the engine centerlineaxis 20.

For most of the axial travel of the first profile 52 and second profile54 along the engine centerline axis 20, the radial distance the firstprofile 52 is disposed from the engine centerline axis 20 differs fromthe radial distance the second profile 54 is disposed from the enginecenterline axis 20. Thus, the volume and cross sectional area of theinlet 18 above the engine centerline axis 20 generally differs from thevolume and cross sectional area of the inlet 18 below the enginecenterline axis 20. The difference in geometry between the first andsecond profiles 52 and 54 is approximated by the inlet centerline axis56 (which approximates the geometric center of the inlet 18), which hasa cambered shape with respect to the engine centerline axis 20.

The inflection point 53 of the first profile 52 is radially and axiallyoffset along the engine centerline axis 20 from the inflection point 55of the second profile 54. More particularly, the convex forward section54F (when viewed from the engine centerline axis 20) of the secondprofile 54 extends further rearward toward the fan face plane 50 thandoes the convex forward section 52F of the first profile 52. Terms suchas “forward,” “rearward,” “upstream,” and “downstream” are defined bythe direction of airflow 22 within the inlet 18. Thus, the concaverearward section 52R (when viewed from the engine centerline axis 20) ofthe first profile 52 extends further forward away from the fan faceplane 50 than does the concave rearward section 54R of the secondprofile 54. This geometry disposes the inflection point 55 between theconvex and concave sections (54F and 54R) of the second profile 54closer to the face of the fan 28 than the inflection point 53 betweenthe convex and concave sections (52F and 54R) of the first profile 52.

In particular, the convex forward section 54F of the second profile 54is geometrically accentuated relative to the concave rearward section54R of the second profile 54 and the convex forward section 52F of thefirst profile 52. Similarly, the concave rearward section 52R of thefirst profile 52 is geometrically accentuated relative to the convexforward section 52F of the first profile 52 and the concave rearwardsection 54R or the second profile 54. The precise interrelation ofprofiles 52 and 54 (i.e. the precise geometric accentuation of theconcave/convex sections 52F, 52R, 54F, 54R relative to one another) isachieved through computational fluid mechanics. The different axial andradial geometry of the inner surface 46 above the engine centerline axis20 with respect to the axial and radial geometry of the inner surface 46below the engine centerline axis 20 cambers the inlet centerline axis 56with respect to the engine centerline axis 20.

In one embodiment, the geometry of the first profile 52 with respect tothe second profile 54 disposes the inlet centerline axis 56 radiallyabove the engine centerline axis 20 immediately to the rear of theforward lip 40 within the inlet 18. The inlet centerline axis 56 extendsabove the engine centerline axis 20 for a portion of its axial extentwithin the inlet 18. In one embodiment, the geometry of the firstprofile 52 with respect to the second profile 54 about the enginecenterline axis 20 causes the inlet centerline axis 56 to have an acutedescending depression angle α at the fan case flange plane 48. In oneembodiment, the acute depression angle α the inlet centerline axis 56forms with respect to the engine centerline axis 20 offsets any residualupwash of the airflow 22 entering the inlet 18.

The geometry of the inner surface 46 allows the airflow 22 in the inlet18 to substantially align with the engine centerline axis 20 (which isalso the rotational axis of the fan 28) at the face of the fan 28. Inthe embodiment shown, the inlet centerline axis 56 coincident with theengine centerline axis 20 at the fan case flange plane 48. In otherembodiments, the inlet centerline axis 56 can merge with or becoincident with the engine centerline axis 20 forward (as defined by thedirection of airflow 22) of the fan case flange plane 48.

As illustrated in FIG. 4, an airflow 22 swirl velocity on the face ofthe fan 28 at 90° and 270° to top dead center (TDC) is reduced oreliminated by offsetting the first profile 52 with respect to the secondprofile 54. The elimination of the swirl velocity occurs because theinlet centerline axis 56 extends below the engine centerline axis 20 andaligns the airflow 22 in the inlet 18 with the engine centerline axis 20at the face of the fan 28. By aligning the airflow 22 in the inlet 18generally with the engine centerline axis 20, the vertical vector of theairflow 22 that causes the airflow 22 swirl velocity is reduced oreliminated. Thus, the swirl velocity at 270° (measured from TDC of thefan 28), co-rotating with the fan 28 (when the fan 28 is rotating in aclockwise direction), and the swirl velocity at 90° counter-rotatingwith the fan 28 (when the fan 28 is rotating in a clockwise direction)are reduced or eliminated. The reduction or elimination of the airflow22 swirl velocity allows the components of the engine 12 to better meetfan stall margin requirements and improves aerodynamic loading.

Although the present invention has been described with reference topreferred embodiments, workers skilled in the art will recognize thatchanges may be made in form and detail without departing from the spiritand scope of the invention.

1. An inlet of a nacelle for channeling inlet airflow to a fan of a gasturbine engine, the inlet comprising: a forward lip extending around acenterline axis of the engine and adapted to accommodate a flow angle ofincoming airflow; an inner surface extending from the forward lip alongthe centerline axis of the engine to adjacent the fan and having a firstprofile above the centerline axis of the engine and a second profilebelow the centerline axis of the engine, the first profile and thesecond profile adapted to define an inlet centerline axis that extendsbelow the centerline axis of the engine at a face of the fan.
 2. Theinlet of claim 1, wherein the inlet centerline axis is cambered withrespect to the centerline axis of the engine to pass above thecenterline axis of the engine for at least a portion of the inlet. 3.The inlet of claim 2, wherein the inlet centerline axis descendsrelative to the centerline axis of the engine and merges with thecenterline axis of the engine substantially at or prior to a plane thatextends along a forward portion of a fan case flange.
 4. The inlet ofclaim 1, wherein the inlet centerline axis forms an acute depressionangle with respect to the centerline axis of the engine, the depressionangle offsets any residual upwash in the inlet airflow such that theairflow is substantially aligned with the centerline axis of the engineat the face of the fan.
 5. The inlet of claim 1, wherein the firstprofile has a concave section that is geometrically accentuated relativeto a convex section of the first profile and a corresponding concavesection of the second profile.
 6. The inlet of claim 5, wherein thesecond profile has a convex section that is geometrically accentuatedrelative to a concave section of the second profile and a correspondingconvex section of the first profile.
 7. The inlet of claim 6, whereinthe degree of geometric accentuation of the first profile sections andthe second profile sections are determined based on computational fluidmechanics.
 8. The inlet of claim 2, wherein the axial and radialgeometry of the inner surface above the engine centerline axis differsfrom the axial and radial geometry of the inner surface below the enginecenterline axis to give the inlet centerline axis the cambered shapewith respect to the engine centerline axis.
 9. An inlet of a nacelle forchanneling inlet airflow to a fan of a gas turbine engine, the inletcomprising: a forward lip extending around a centerline axis of theengine; a crown portion defining the forward lip substantially above thecenterline axis of the engine and having an inner surface extending fromthe forward lip to adjacent the fan, at least a portion of the innersurface has a concave shape; and a keel portion defining the forward lipsubstantially below the centerline axis of the engine, the keel portionhaving an inner surface that extends from the forward lip to adjacentthe fan, at least a portion of the inner surface has a convex shape;wherein the inner surface of the crown portion and the inner surface ofthe keel portion define a cambered inlet centerline axis that forms adescending depression angle with the centerline axis of the engine atthe fan case flange.
 10. The inlet of claim 9, wherein the depressionangle defined by the inner surface of the crown portion and the innersurface of the keel portion offsets residual upwash in the inlet airflowsuch that the airflow is substantially aligned with the centerline axisof the engine at the face of the fan.
 11. The inlet of claim 9, whereinthe cambered centerline axis defined by the inner surface of the crownportion and keel portion passes above the centerline axis of the enginefor at least a portion of the inlet.
 12. The inlet of claim 11, whereinthe inlet centerline axis defined by the inner surface of the crownportion and keel portion merges the inlet centerline axis with thecenterline axis of the engine at or upstream of a plane that extendsalong a forward portion of a fan case flange.
 13. The inlet of claim 12,wherein the inlet centerline axis descends relative to the centerlineaxis of the engine and merges with the centerline axis of the enginesubstantially at the plane that extends along the forward portion of thefan case flange.
 14. The inlet of claim 9, wherein the inner surface ofthe crown portion has a concave section which extends further forwardaway from the face of the fan than a corresponding concave section ofthe inner surface of the keel portion.
 15. The inlet of claim 14,wherein the inner surface of the keel portion has a convex section whichextends further rearward toward the face of the fan than a correspondingconvex section of the inner surface of the crown portion.
 16. The inletof claim 15, wherein an inflection point between the convex and concavesections of the inner surface of the keel portion occurs closer to theface of the fan than an inflection point between the convex and concavesections of the inner surface of the crown portion.
 17. The inlet ofclaim 9, wherein the volume of the inlet above the centerline axis ofthe engine adjacent the fan exceeds the volume of the inlet below thecenterline axis of the engine adjacent the fan.